Most of the components and subsystems of a spacecraft must operate in restricted temperature ranges. This makes thermal control a key matter in the design and operation of a spacecraft with a significant weight, power and cost impact in the overall spacecraft budgets.
Spacecraft thermal control relies on the global spacecraft thermal balance: the heat loads must be rejected to deep space that works as a thermal sink. Since no matter links this sink and the spacecraft, this rejection is made by thermal radiation through dedicated radiators installed on the satellite external surfaces.
Spacecraft thermal loads come from the internal spacecraft equipment dissipation and, externally, from the sun and the earth or from the celestial bodies around which the spacecraft orbits. The thermal systems used in spacecrafts must therefore be able to control equipment which operates at a very high temperature and also discontinuously.
Current thermal control state of the art is based on passive and active methods, these methods depending on elements requiring or not power to be functional. Some of these known elements are coatings, Multi Layer Insulation (MLI), heaters, heat pipes, Loop Heat Pipes, Capillary Pumped Loops, Mechanically Pumped Loops, etc. with insulation, radiation, heat transportation, temperature homogenisation or heating functions. Given the variety of thermal requirements and the harsh space environment, these thermal elements must be selected, designed, manufactured and integrated very carefully.
Document U.S. Pat. No. 4,162,701 discloses a thermal control canister for a spacecraft, maintained at a substantially constant temperature. Fixed conductance heat pipes on the canister walls are connected to variable conductance heat pipes (VCHP), mounted on the radiator structure. The effective radiating area of the radiator structure is controlled by the VCHP in response to sensed temperature of the instrument package or the canister wall. This comparison controls a heater in a gas reservoir containing a non-condensable gas of the VCHP. The VCHP can either be located between the canister and radiators or can be coupled directly between the canister walls and one or more radiators. This solution can be applied on element level but it is difficult to be used for the thermal control of an entire spacecraft. This design is very heavy and very expensive for small spacecrafts. Moreover, additional special systems will be required for large satellites to collect and transfer heat from onboard equipment and to distribute this heat on radiators with VCHP. This makes this design very complex, not very efficient (many thermal interfaces) and not reliable. Also, VCHP are not flexible enough and capable to transfer high power (maximum several hundreds of watts) for shorter distances (up to 2-3 m).
Document U.S. Pat. No. 6,478,258, discloses a loop heat pipe for use on a spacecraft. The loop heat pipe cooling system comprises loop heat pipes routed from internally facing surfaces of one or more internally located equipment panels to externally located radiator panels. Heat is collected at evaporator ends of each loop heat pipe and is transported to condenser ends of the respective loop heat pipe. The loop heat pipes used in the cooling system are flexible and easily routed, so that they can be routed to multiple radiator panels in order to optimize heat sharing between radiator panels. The total number of loop heat pipes used in the cooling system depends on the overall heat load. The system also comprises one or more fixed conductance heat pipes mounted to selected internally facing surfaces of the internally located equipment panels. The problem of this system is an impossibility to control the temperature of the equipment since loop heat pipes are just heat transfer devices.
Document JP 2001315700 discloses a thermal control system for a spacecraft, the system minimizing the generation of vibration and inertia force by eliminating or minimizing rotation of a radiator for radiating heat into space. The system comprises a radiator panel, a control unit and selector valves, the heat generated inside the spacecraft being radiated into space by switching the selector valves without rotation of the radiator panel, so that generation of inertia force or vibration is prevented. The problem of this system is connected with thermal design complexity: two opposite radiators have to be well thermally disconnected but it is difficult to reach this aim since one radiator with embedded heat exchanger is placed on the top of another. Therefore, the thermal control system of JP 2001315700 is not capable of providing good temperature stabilization in a narrow range (several degrees).
Document U.S. Pat. No. 6,073,888, which is considered as the closest prior art of the invention, and upon which the preamble of claim 1 is based, discloses an increased satellite heat rejection system comprising radiating surfaces which are exposed to direct sun light on an intermittent basis. The system is applied to earth-orbiting satellites, especially to those in a geosynchronous orbit, the system comprising a thermal radiator mounted on a face for discharging heat from a thermal load to deep space. A heat conductor extends between the thermal load and the thermal radiator. The system also comprises thermal switches operable for connecting the thermal load to the thermal radiator for cooling when the temperature of the thermal load is above a predetermined level and for disconnecting the thermal load from the thermal radiator when the temperature of the thermal load falls below the predetermined level. This invention is based on VCHP architecture with active temperature control. Heaters are installed on VCHP reservoirs filled by non-condensable gas. Computer governs the heater power as a function of radiators temperatures (dedicated temperature sensors have to be installed on every radiator). The disadvantages of VCHP were already discussed above. Also, although this system is a passive thermal control system but heaters and control electronics require devoted power budget, which is a critical issue for space applications. The thermal switches in U.S. Pat. No. 6,073,888 are operating as ON/OFF devices: such type of control is not precise and sensitive enough for thermal systems, where thermal inertia plays an important role.
The present invention is oriented to the solution of the above-mentioned drawbacks.